1. Field of the Invention
This invention relates to a supersonic vortex generator for structures which are subjected to supersonic airflow that normally results in flow separation, wave drag and other interference with airflow, to attenuate that flow separation, induced drag and airflow interruption.
In particular, the invention concerns vortex generators comprising a series of triangular cavities in an airfoil or other surfaces over which supersonic air is flowing, which are capable of generating a series of initially diverging and then streamwise spiral vortices trailing after the cavities that serve to attenuate flow separation and wave drag penalties.
2. Description of the Prior Art
Flow separation in subsonic airflow over structures such as aircraft airfoils is prevented or lessened by vortex generators that project from the surface of the wing or other structure. The vortex generators primarily consist of a plurality of short, low-aspect-ratio airfoils arranged in pairs extending away from the surface of the airfoil.
In a typical swept wing subsonic aircraft wing application, the vortex generators may comprise a series of relatively short, upstanding blades arranged such that adjacent pairs include one blade that is generally aligned with the path of air over the wing, while an adjacent blade may be at an acute angle of about 20.degree. with respect to the line of flight. The tip vortices of these blade airfoils pull high-energy air down into the boundary layer of the wing and prevent flow separation. Blade vortex generators do not work though in supersonic vehicle applications because of the excessive wave drag penalty that results.
When air at supersonic speed flows over a corner surface that is concave, the flow must remain tangent to the surfaces; hence, the streamline at the corner is deflected to conform to the oblique angle of the turned surface. Whenever a supersonic flow is turned "into itself" an oblique shock wave occurs A similar phenomenon happens when supersonic flow passes over a wedge-shaped object, or a cone-shaped object. Across the shock wave, the Mach number decreases, but the pressure, the temperature and the density of the air increase. The impact with the ground of strong shock waves formed on aircraft during supersonic flight causes a loud sound called a "sonic boom".
Turning of a supersonic flow "away from itself", causes an expansion fan to occur. This expansion wave is in the shape of a fan centered at the corner of the two surfaces which are oblique to one another. The fan continuously opens in a direction away from the corner. The originally horizontal streamline flows ahead of the expansion wave are deflected smoothly and continuously through the expansion wave causing the streamlines behind the wave to be parallel to each other and inclined downwardly at the deflection angle of one surface with respect to the other surface. Across the expansion wave, the Mach number increases, but the pressure, the temperature and the density of the air decrease.
In actual supersonic flight of an aircraft, a combination of shock waves and expansion fans form on different parts of the vehicle depending on their shape and location. Typically, strong shock waves occur at the nose of the aircraft and at the middle section, with expansion fans being formed in the aft body area.
An aircraft wing stalls whenever the smooth airflow over its top surface separates to create a turbulent region over the wing surface. At subsonic speeds, stalls occur at high angle of attack flight conditions because the airflow tends to separate from the top surface of the airfoil, creating a large wake of relatively "dead air" behind the airfoil. Inside of this separated region, the flow is recirculating, and part of the flow actually moves in a direction opposite to the free stream, creating a so-called "reversed flow". The consequence of this separated flow at a high angle of attack is a sudden decrease in lift and a large increase in drag causing a stall.
At supersonic speeds, a stall can occur as a result of shock-induced separation of airflow across the wing or other part of the aircraft, regardless of the angle of attack.
Mach number is the ratio of the aircraft speed to the speed of sound. When a vehicle reaches its critical Mach number, i.e., when it flies at a speed at which the airflow at any portion of its surfaces equals the speed of sound, a shock wave begins to form just behind the point at which the air is moving the fastest. This shock wave oscillates back and forth and causes the air to separate from the upper surface of the airfoil. This flow separation results in loss of lift and can ultimately cause a stall. There is also a concomitant increase in drag and buffeting of control surfaces attached to the trailing edge of the wing. Similar separation conditions can occur on jet propulsion engine intakes, compressor blades, jet engine exhaust nozzles and many other areas of the vehicle.